CN108931791B - System and method for correcting satellite inertial force combined clock difference - Google Patents
System and method for correcting satellite inertial force combined clock difference Download PDFInfo
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01S—RADIO DIRECTION-FINDING; RADIO NAVIGATION; DETERMINING DISTANCE OR VELOCITY BY USE OF RADIO WAVES; LOCATING OR PRESENCE-DETECTING BY USE OF THE REFLECTION OR RERADIATION OF RADIO WAVES; ANALOGOUS ARRANGEMENTS USING OTHER WAVES
- G01S19/00—Satellite radio beacon positioning systems; Determining position, velocity or attitude using signals transmitted by such systems
- G01S19/01—Satellite radio beacon positioning systems transmitting time-stamped messages, e.g. GPS [Global Positioning System], GLONASS [Global Orbiting Navigation Satellite System] or GALILEO
- G01S19/13—Receivers
- G01S19/23—Testing, monitoring, correcting or calibrating of receiver elements
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
- G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
- G01C21/10—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration
- G01C21/12—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning
- G01C21/16—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation
- G01C21/165—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation combined with non-inertial navigation instruments
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
- G01C25/00—Manufacturing, calibrating, cleaning, or repairing instruments or devices referred to in the other groups of this subclass
- G01C25/005—Manufacturing, calibrating, cleaning, or repairing instruments or devices referred to in the other groups of this subclass initial alignment, calibration or starting-up of inertial devices
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01S—RADIO DIRECTION-FINDING; RADIO NAVIGATION; DETERMINING DISTANCE OR VELOCITY BY USE OF RADIO WAVES; LOCATING OR PRESENCE-DETECTING BY USE OF THE REFLECTION OR RERADIATION OF RADIO WAVES; ANALOGOUS ARRANGEMENTS USING OTHER WAVES
- G01S19/00—Satellite radio beacon positioning systems; Determining position, velocity or attitude using signals transmitted by such systems
- G01S19/01—Satellite radio beacon positioning systems transmitting time-stamped messages, e.g. GPS [Global Positioning System], GLONASS [Global Orbiting Navigation Satellite System] or GALILEO
- G01S19/13—Receivers
- G01S19/24—Acquisition or tracking or demodulation of signals transmitted by the system
- G01S19/25—Acquisition or tracking or demodulation of signals transmitted by the system involving aiding data received from a cooperating element, e.g. assisted GPS
- G01S19/256—Acquisition or tracking or demodulation of signals transmitted by the system involving aiding data received from a cooperating element, e.g. assisted GPS relating to timing, e.g. time of week, code phase, timing offset
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01S—RADIO DIRECTION-FINDING; RADIO NAVIGATION; DETERMINING DISTANCE OR VELOCITY BY USE OF RADIO WAVES; LOCATING OR PRESENCE-DETECTING BY USE OF THE REFLECTION OR RERADIATION OF RADIO WAVES; ANALOGOUS ARRANGEMENTS USING OTHER WAVES
- G01S19/00—Satellite radio beacon positioning systems; Determining position, velocity or attitude using signals transmitted by such systems
- G01S19/38—Determining a navigation solution using signals transmitted by a satellite radio beacon positioning system
- G01S19/39—Determining a navigation solution using signals transmitted by a satellite radio beacon positioning system the satellite radio beacon positioning system transmitting time-stamped messages, e.g. GPS [Global Positioning System], GLONASS [Global Orbiting Navigation Satellite System] or GALILEO
- G01S19/42—Determining position
- G01S19/45—Determining position by combining measurements of signals from the satellite radio beacon positioning system with a supplementary measurement
- G01S19/47—Determining position by combining measurements of signals from the satellite radio beacon positioning system with a supplementary measurement the supplementary measurement being an inertial measurement, e.g. tightly coupled inertial
Abstract
The invention relates to a satellite inertial compact combined clock difference correction system and a method, which comprises a satellite navigation receiver, an inertial navigation system and a satellite inertial compact combined clock difference correction device, wherein the satellite navigation receiver is connected with the inertial navigation system and the satellite inertial compact combined clock difference correction device, the inertial navigation system is connected with the satellite inertial compact combined clock difference correction device, the satellite navigation receiver carries out processing and PVT calculation according to received satellite signals to realize positioning of the satellite navigation receiver and adjust the clock difference of the receiver, the inertial navigation system carries out strapdown calculation to obtain the position, the speed and the attitude information of the inertial navigation system, calculates a pseudo range estimation value and a pseudo range rate estimation value according to the received data, the satellite inertial navigation system carries out compact combined clock difference correction on the position, the speed and the attitude information of the inertial navigation system according to the received data to obtain and output the corrected position, speed and attitude information, and is compatible with positioning and satellite inertial compact combined positioning, the clock error of the receiver is used for correcting the clock error of the tight combination, so that the navigation accuracy is effectively improved.
Description
Technical Field
The invention relates to the field of integrated navigation, in particular to a satellite inertial tightening integrated clock difference correction system and a satellite inertial tightening integrated clock difference correction method.
Background
The satellite signals are tracked by the satellite navigation receiver to generate two basic distance measurement values of pseudo range and carrier phase. The measurement of pseudoranges is closely related to time. The time generated by the receiver clock is usually not synchronous with the time of the satellite guide, the time difference between the two is usually called as the receiver clock difference, and in the positioning process, the clock difference value needs to be estimated to finish the accurate time service of the system to the outside. The clock offset is affected by the clock accuracy and drifts over time. When the clock error is not corrected, the system time is inaccurate, so that the locally copied signal is not synchronous with the signal received by the receiver, and even the signal is out of lock. Therefore, the system needs to adjust the clock difference.
In a traditional tight combination algorithm, a receiver and a tight combination do not work simultaneously, a positioning result of the receiver is usually completely replaced by a positioning result of the tight combination, and the tight combination is matched with the adjustment of the clock error of the receiver, so that the real-time error condition of the system is difficult to accurately reflect, the observation of the tight combination is influenced, and the navigation positioning accuracy is low.
Disclosure of Invention
In view of the above, there is a need to provide a system and a method for correcting a satellite inertial force difference with high navigation positioning accuracy.
A satellite inertial compression combined clock difference correction system comprises a satellite guide receiver, an inertial guide system and a satellite inertial compression combined clock difference correction device, wherein the satellite guide receiver is connected with the inertial guide system and the satellite inertial compression combined clock difference correction device, the inertial guide system is connected with the satellite inertial compression combined clock difference correction device,
the satellite guide receiver is used for receiving satellite signals and processing the satellite signals to obtain pseudo ranges, pseudo range rates and positions and speeds of satellites between the satellites and the satellite guide receiver; performing PVT (physical vapor transport) calculation according to the pseudo range and the pseudo range rate to obtain the position, the speed, the clock error and the clock drift of the guide receiver; adjusting the current clock difference according to the clock difference and the clock drift to obtain an actual adjusted clock difference, sending the pseudo range, the pseudo range rate and the actual adjusted clock difference to the inertial navigation combined clock difference correction device, and sending the position and the speed of the satellite and the position and the speed of the inertial navigation receiver to the inertial navigation system;
the inertial navigation system is used for obtaining position, speed and attitude information of the inertial navigation system through strapdown resolving, calculating to obtain a pseudo-range estimation value and a pseudo-range rate estimation value by combining the received position and speed of the satellite, and sending the pseudo-range estimation value, the pseudo-range rate estimation value and the position, speed and attitude information of the inertial navigation system to the satellite inertial navigation tight combination clock error correction device;
and the satellite-inertial tight combination clock difference correction device is used for performing tight combination clock difference correction on the position, speed and attitude information of the inertial navigation system according to the received pseudo range, the pseudo range rate, the actual adjustment clock difference, the pseudo range estimation value and the pseudo range rate estimation value to obtain and output the position, speed and attitude information after the inertial navigation system is corrected.
A satellite inertial tightening combined clock difference correction method comprises the following steps:
the satellite signal is received by the satellite guide receiver and processed to obtain a pseudo range, a pseudo range rate and a position and a speed of the satellite between the satellite and the satellite guide receiver;
the satellite navigation receiver carries out PVT resolving according to the pseudo range and the pseudo range rate to obtain the position, the speed, the clock error and the clock drift of the satellite navigation receiver, and the current clock error is adjusted according to the clock error and the clock drift to obtain an actual adjusted clock error;
the satellite navigation receiver sends the pseudo range, the pseudo range rate and the actual adjustment clock difference to the satellite inertial tight combination clock difference correction device, and sends the position and the speed of the satellite and the position and the speed of the satellite navigation receiver to the inertial navigation system;
the inertial navigation system obtains position, speed and attitude information of the inertial navigation system through strapdown resolving, and calculates a pseudo-range estimation value and a pseudo-range rate estimation value by combining the received position and speed of the satellite;
the inertial navigation system sends the pseudo-range estimated value, the pseudo-range rate estimated value and position, speed and attitude information of the inertial navigation system to the satellite-inertial tight combination clock difference correction device;
and the satellite-inertial tight combination clock difference correction device performs tight combination clock difference correction on the position, speed and attitude information of the inertial navigation system according to the received pseudo range, the pseudo range rate, the actual adjustment clock difference, the pseudo range estimation value and the pseudo range rate estimation value, so as to obtain and output the position, speed and attitude information corrected by the inertial navigation system.
The satellite inertial tight combined clock difference correction system and method are compatible with positioning of the satellite navigation receiver and positioning of the satellite inertial tight combined, correct the tight combined clock difference through the receiver clock difference, adjust the clock difference in real time during tight combination, and effectively improve navigation accuracy.
Drawings
FIG. 1 is a block diagram of a tight combination clock correction system in accordance with an embodiment;
FIG. 2 is a flow chart of a method for tightly combining clock offset corrections in an embodiment;
FIG. 3 is a flow chart of a method for correcting tightly combined clock skew in another embodiment.
Detailed Description
In one embodiment, as shown in fig. 1, a satellite inertial navigation combined clock difference correction system includes a satellite navigation receiver 110, an inertial navigation system 120, and a satellite inertial navigation combined clock difference correction device 130, where the satellite navigation receiver 110 is connected to the inertial navigation system 120 and the satellite inertial navigation combined clock difference correction device 130, the inertial navigation system 120 is connected to the satellite inertial navigation combined clock difference correction device 130, and the satellite navigation receiver 110 is configured to receive satellite signals and process the satellite signals to obtain pseudoranges, pseudorange rates, and positions and velocities of the satellites between the satellites and the satellite navigation receiver 110; performing PVT (Position, Velocity, Time) solution according to the pseudo-range and the pseudo-range rate to obtain a Position, a Velocity, a clock difference, and a clock drift of the guide receiver 110; adjusting the current clock difference according to the clock difference and the clock drift to obtain an actual adjusted clock difference, sending the pseudo-range, the pseudo-range rate and the actual adjusted clock difference to the inertial navigation combination clock difference correction device 130, and sending the position and the speed of the satellite and the position and the speed of the inertial navigation receiver 110 to the inertial navigation system 120; the inertial navigation system 120 is configured to obtain position, velocity, and attitude information of the inertial navigation system 120 through strapdown solution, and calculate a pseudorange estimation value and a pseudorange rate estimation value in combination with the received position and velocity of the satellite; the inertial navigation system 120 sends the pseudo-range estimation value, the pseudo-range rate estimation value and the position, speed and attitude information of the inertial navigation system 120 to the inertial tight combination clock difference correction device 130; the inertial tight combination clock difference correction device 130 is configured to perform tight combination clock difference correction on the position, speed, and attitude information of the inertial navigation system 120 according to the received pseudorange, pseudorange rate, actual adjustment clock difference, pseudorange estimation value, and pseudorange rate estimation value, to obtain and output position, speed, and attitude information after being corrected by the inertial navigation system 120.
In particular, the satellite receiver 110, i.e., the GPS receiver, is an instrument that receives global positioning system satellite signals and determines the ground spatial location. The navigation positioning signal sent by the GPS satellite is an information resource which can be shared by countless users, and for vast users on land, sea and space, the GPS signal receiver only needs to be provided with a receiving device capable of receiving, tracking, transforming and measuring the GPS signal. The inertial navigation system is an autonomous navigation system which does not depend on external information and radiates energy to the outside. The working environment of the device not only comprises the air and the ground, but also can be underwater. The basic working principle of the inertial navigation system is based on Newton's law of mechanics, and by measuring the acceleration of a carrier in an inertial reference system, integrating the acceleration with time and transforming the acceleration into a navigation coordinate system, information such as speed, yaw angle, position and the like in the navigation coordinate system can be obtained; the close combination realizes the combination of the satellite navigation and the inertial navigation on a measurement domain, namely, the pseudorange and the pseudorange rate are taken as the measurement of the system, so as to realize the satellite-inertial-close-combination (inertial satellite-close-combination) positioning. The tight combination is relatively loose combination, and the precision is higher, the interference killing feature is stronger, for dark combination, realizes simply, and the degree of difficulty is little. However, because pseudo-range is used as observation, the tight package of samples requires estimation and adjustment of the system clock error.
Specifically, the satellite receiver 110 receives a satellite signal from an antenna, and processes the satellite signal to obtain a pseudorange and a pseudorange rate between the satellite and the satellite receiver 110, which includes the following steps: down-converting the guard pilot radio frequency signal into an intermediate frequency signal, wherein an intermediate frequency point is a preset frequency; after acquiring, tracking, bit synchronizing, frame synchronizing and text demodulating the satellite pilot intermediate frequency signal, calculating to obtain a pseudo range and a pseudo range rate between the satellite and the satellite pilot receiver 110, specifically:
where ρ is(n)、r(n)、δt(n)、I(n)、T(n)、
δtuRespectively a pseudo-range measurement value corresponding to a satellite n, a distance measurement value corresponding to a receiver geometric distance, a satellite clock error and an ionosphere error, a distance measurement value corresponding to a troposphere error, a pseudo-range measurement error and a receiver clock error;
respectively correspond to rho(n)、r(n)、δt(n)、I(n)、T(n)、
δtuThe derivative of (a) is a pseudo-range rate measurement value corresponding to the satellite n, a range change rate measurement value corresponding to the receiver, a satellite clock speed, a range change rate measurement value corresponding to the ionosphere error, a range change rate measurement value corresponding to the troposphere error, a pseudo-range rate measurement error, and a receiver clock speed.
Specifically, the position and velocity of the satellites are resolved from the received satellite ephemeris.
In one embodiment, after the satellite signal is processed to obtain the pseudorange, the pseudorange rate, and the position and the velocity of the satellite, the satellite receiver 110 performs PVT solution according to the pseudorange and the pseudorange rate to obtain the position, the velocity, the clock offset, and the clock drift of the satellite receiver 110, which includes: the satellite navigation receiver is used for acquiring the number of connected satellites, and when the number of the satellites is larger than or equal to a preset threshold value, PVT resolving is carried out according to the pseudo range and the pseudo range rate to obtain the position, the speed, the clock difference and the clock drift of the satellite navigation receiver; and when the number of the satellites is smaller than the preset threshold value, returning to receive the satellite signals, and processing the satellite signals to obtain the pseudo range, the pseudo range rate, the position and the speed of the satellites.
Specifically, the pseudorange and the pseudorange rate between the satellite and the satellite guide receiver 110 are used as observed quantities, PVT calculation is performed according to the relationships between the observed quantities and the position, the speed, the clock difference and the clock drift of the satellite guide receiver 110, the PVT calculation refers to the position, the speed and the time calculation of the user receiver as the name implies, that is, when the number of satellites is greater than or equal to 4, the position, the speed, the clock difference and the clock drift of the satellite guide receiver 110 are calculated by using a (weighted) least square method or a filtering method, and a specific formula is as follows, and a receiver positioning result can be output as a system result:
wherein x ═ x y z]TFor unknown receiver position coordinate vectors, x(n)＝[x(n) y(n) z(n)]TIs the position coordinate vector of the satellite n,
the error amount of pseudo-range measurement is omitted in the formula for the pseudo-range measurement value after error correction
Obtaining an equation set of solving speed after derivation on two sides of the above formula, wherein the equation set is not given; clock error is delta tuThe clock drift resolves the clock speed inside for the speed.
And correspondingly advancing or delaying the local time according to the clock difference and the clock drift obtained by calculation, namely obtaining the actual adjustment clock difference for adjusting the current clock difference of the receiver, so as to generate a corresponding local pseudo-satellite signal and finish the continuous tracking of the satellite signal.
In one embodiment, the inertial navigation system 120 is configured to obtain position, velocity, and attitude information of the inertial navigation system through strapdown solution, and obtain a pseudorange estimation value and a pseudorange rate estimation value by combining the position and velocity of the received satellite, specifically:
pseudo range estimation value-satellite position-inertial navigation system position
Pseudo range rate estimation value being satellite velocity-inertial navigation system velocity
Specifically, the inertial navigation system 120 obtains the position, velocity, and attitude information of the inertial navigation system 120 through strapdown calculation, where the strapdown calculation performed by the inertial navigation system 120 is mechanical arrangement of the inertial navigation, including position update, velocity update, and attitude update of the inertial navigation, that is, the position, velocity, and attitude information of the inertial navigation system 120 are obtained by numerical integration (using, for example, a quaternion method, an euler angle method, a direction cosine method, and the like) using an angular velocity output by a gyroscope and acceleration information output by an accelerometer, and a specific formula is shown below, where the inertial navigation system 120 is mounted on a carrier, and the position, velocity, and attitude information of the inertial navigation system 120 can also be understood as positioning of the carrier, and can also be output as a system result.
a) Attitude updating of inertial navigation:
the transformation matrix between the navigation coordinate system and the body coordinate system is
Is provided with
Then:
θ＝-sin-1(C31)
wherein, the pitch angle θ: (-90, 90) course angle
(-180, 180), roll angle γ: (0, 360).
The attitude differential equation corresponding to the attitude matrix is:
wherein the content of the first and second substances,
a coordinate transformation matrix for the navigation coordinate system to the carrier coordinate system,
is a derivation of the matrix and,
to correspond to the angular velocity of attitude
Forming an anti-symmetric matrix.
The attitude angular rate is:
wherein the content of the first and second substances,
is the output of the angular rate gyroscope,
is the angular rate of rotation of the earth,
to navigate the angular velocity of the coordinate system relative to the earth, the instantaneous velocity can be determined
And finding out the position.
b) And (3) inertial navigation speed updating:
the velocity update of inertial navigation is obtained by solving the inertial navigation velocity differential equation as follows:
wherein, Ve、Vn、VuRespectively velocity in the northeast direction, omegaieIs the angular velocity of rotation, f, of the terrestrial coordinate system relative to the inertial coordinate systeme、fn、fuRespectively, specific force information R output by the accelerometer in the northeast directionNIs the curvature radius of the mortise-tenon unitary ring, RMIs the radius of curvature of the meridian,
h is the geographical altitude and g is the acceleration of gravity.
c) Position updating of inertial navigation:
after solving for velocity, the position can be obtained by velocity integration:
where λ, L, h are the receiver longitude, latitude, and elevation, respectively.
In one embodiment, the device for correcting satellite inertial combined clock difference 130 is configured to perform close combined clock difference correction on the position, speed and attitude information of the inertial navigation system 120 according to the received pseudorange, pseudorange rate, actual adjusted clock difference, pseudorange estimation value and pseudorange rate estimation value, obtain and output the corrected position, speed and attitude information of the inertial navigation system 120, and includes: the satellite-inertial tight combination clock error correction device 130 is used for performing tight combination filtering according to the pseudo range, the pseudo range rate, the actual adjustment clock error, the pseudo range estimation value and the pseudo range rate estimation value, and calculating to obtain an error of the inertial navigation system; the satellite-inertial tight combination clock difference correction device 130 is configured to correct the position, speed, and attitude information of the inertial navigation system according to the error of the inertial navigation system, obtain the corrected position, speed, and attitude information of the inertial navigation system, and output the position, speed, and attitude information.
Specifically, in the satellite-inertial navigation part, a Kalman filtering algorithm is adopted, and the Kalman filtering algorithm comprises a time updating process and an observation updating process. The time updating process comprises one-step prediction of state quantity and mean square error, and the observation updating process comprises state estimation and mean square error estimation, and specifically comprises the following steps:
a) one-step prediction equation of system state
In the formula (I), the compound is shown in the specification,
is in a state Xk-1Kalman filter estimation, using tk-1The observed values of the time and the previous time are calculated,
is to utilize
Calculated pair XkThe one-step prediction of (1) can also be considered as using tk-1The observed value of time and previous time is tkOne step prediction of time of day,. phik,k-1Is a state transition matrix.
b) System state estimation equation
The above formula is calculated based on one-step prediction according to the observed value,
can be called an innovation process, ZkTo observe the vector, HkTo observe the transition matrix.
c) Filter gain equation
Wherein, KkGain matrix, R, in Kalman Filter Algorithm iterative ProcesskError covariance matrix, K, of observed noisekThe criterion chosen is to minimize the mean square error matrix of the estimated quantities. RkBig rule KkSmall, the estimated quantity has small dependence degree on the observed value, RkMinor rule KkAnd large, the estimated quantity has large dependence on the observed value.
d) One-step prediction mean square error equation
Γk-1For the system noise matrix, a one-step prediction mean square error must be solved before the filter gain matrix is solved. One-step prediction mean square error array Pk/k-1Is in mean square error matrix Pk-1Considering the system noise variance matrix Q on the basis ofk-1The effect of (c) is obtained.
e) Estimating mean square error equation
Specifically, in the present scheme, the tight combination filtering is the kalman filtering, that is, the tight combination filtering implemented by the kalman filtering, where the tight combination filtering refers to the optimal estimation of the kalman filtering, and the kalman filtering is divided into 5 steps: one-step prediction, one-step covariance prediction, a gain matrix, an optimal estimate, and an optimal error covariance estimate.
In the implementation process, a state equation of the system is established according to an error propagation model of inertial navigation and a dynamic model of a clock-error clock drift, the dynamic model of the clock-error clock drift can usually adopt a random walk model or a first-order markov model, and an observation equation is established according to a relation between an observed quantity and a state quantity, which is specifically as follows:
a) tightly packed equation of state
The state quantity is selected as
Wherein, subscripts e, n and u represent the three axial directions of east, north and sky, phie、φn、φuIs platform error angle of east, north and sky, and is position error of longitude, latitude and height, and is Δ Ve、ΔVn、ΔVuFor east, north and sky speed errors, deltatuAnd δ truThe clock difference and the clock speed of the clock of the GPS receiver.
The state quantity is an inertial navigation error parameter, and the state equation is selected as an error propagation equation of inertial navigation:
wherein, XIIs a state vector, AIThe matrix is a state transition matrix, GIFor noise-driven matrix, WIFor system driving noise, each parameter can be obtained by an inertial navigation error propagation equation and a receiver clock error model.
b) Tightly combined observation equation
The close-coupled observation equations include pseudorange error observations and pseudorange rate error observations. The tight combination provides a pseudorange p measured by the receiverGjPseudorange rate
And pseudo range estimated value rho obtained by inertial navigation calculationIjAnd pseudo range rate estimate
Difference rho between pseudo distances obtained by the twoGj-ρIjAnd the difference between pseudo range rates
The pseudo range and pseudo range rate observed quantity are used as the closely combined system.
The observation equation of the pseudo-range error is:
Zρ(t)＝Hρ(t)X(t)+Vρ(t)
in the formula:
Zρ(t)＝δρj＝ρGj-ρIj
Hρ(t)＝[0j×3 aj1 aj2 0j×2 aj3 0 Hρ1]j×11
aj1＝-(RN+h)[ej1cos L sinλ-ej2 cos L cosλ]
aj2＝(RN+h)[-ej1 sin L cosλ-ej2 sin L sinλ]+[RN(1-e2)+h]ej3 cos L
aj3＝ej1 cos L cosλ+ej2 cos L sinλ+ej3 sin L
wherein Z isρ(t) is the pseudorange error observed, ρGj、ρIjRespectively is the jth satellitePseudoranges measured by a satellite-associated receiver and estimated pseudoranges derived from inertial navigation Hρ(t) is a pseudo-range error observation matrix, (x)I yI zI)TFor the position measured by inertial navigation, the jth satellite position determined from the satellite ephemeris is (x)sj ysj zsj)T，
To make the partial derivatives for x, e is the eccentricity of the ellipse, where,
a is the major radius of the reference ellipsoid, b is the minor radius of the reference ellipsoid, rjIs the geometric distance of the receiver from the satellite.
The pseudo-range rate error observation equation is as follows:
in the formula:
bj1＝-ej1 cosλsin L-ej2 sin L sinλ+ej3 cos L
bj2＝-ej1 sinλ+ej2 cosλ
bj3＝ej1 cos L cosλ+ej2 cos L sinλ+ej3 sin L
wherein the content of the first and second substances,
is the pseudorange rate error observation,
respectively are the pseudo range rate measured by the receiver corresponding to the jth satellite and the estimated value of the pseudo range rate calculated by inertial navigation,
is a pseudo-range rate error observation matrix, X (t) is a system state vector, Vp(t) is the observation noise vector.
Combining the pseudo-range measurement equation and the pseudo-range rate measurement equation into a measurement equation of the combined navigation system, wherein the observed quantity consists of pseudo-range difference and pseudo-range difference to form a multi-dimensional observation vector, and the measurement equation of the combined system can be expressed as:
and performing tight combination filtering according to the observation equation to estimate the state quantity.
Generally, the resolving frequency of inertial navigation is higher than that of satellite navigation PVT, a satellite navigation PVT resolving period is used as a system filtering estimation period in the filtering process, and the inertial navigation resolving period is used as a system filtering prediction period. And in each epoch, the pseudo range and the pseudo range rate between the satellite and the satellite receiver 110, which are obtained by tracking by the satellite receiver 110, and the pseudo range estimation value and the pseudo range rate estimation value calculated by inertial navigation are subjected to difference to be used as the observed quantity of the tight combination filtering, specifically, after the tight combination filtering is finished, the error of inertial navigation, including a position error, a speed error and an attitude error, is obtained through estimation, and finally, the final position, the speed and the attitude of the inertial navigation system 120 are obtained through feedback correction.
In one embodiment, the device for correcting the inertial tightly combined clock difference 130 is configured to perform tightly combined filtering according to the pseudorange, the pseudorange rate, the actual adjusted clock difference, the pseudorange estimated value, and the pseudorange rate estimated value, and calculate an error of the inertial navigation system, including: the satellite inertial force combined clock difference correction device 130 is used for obtaining the clock difference of the current moment predicted by the clock difference model; the Wei-inertial fastening combination clock difference correction device 130 is used for calculating a predicted value of the clock difference at the current time according to the clock difference at the current time predicted by the clock difference model and the actually adjusted clock difference; and the satellite-inertial tight combination clock difference correction device 130 is used for performing tight combination filtering according to the clock difference predicted value, the pseudo range rate, the pseudo range estimated value and the pseudo range rate estimated value at the current moment, and calculating to obtain the error of the inertial navigation system.
Specifically, the tight combination filtering process is a tight combination state equation and a measurement equation, and is iterated through 5 steps of kalman filtering. In the filtering process, because the guard receiver 110 itself has a clock difference model, the tight combination models the clock difference, and both models are complete to realize. When the clock difference estimation method is used together, conflicts exist, on the basis of PVT calculation based on a receiver, the receiver adjusts local time by utilizing the clock difference estimated by the receiver, so that the clock difference estimation of the tight combination can not be obtained according to the change rule described by the model, the clock difference of the tight combination needs to be reset in each epoch, namely, the clock difference of the current time is obtained by predicting the tight combination according to a clock difference clock drift model (such as a random walk model or a first-order Markov model), and the clock difference quantity actually adjusted by the receiver of the previous epoch is subtracted, so that the clock difference prediction value of the tight combination current time is obtained. At this time, the troposphere error, the ionosphere error and the clock error predicted value are added to the distance between the satellite and the carrier calculated by inertial navigation, the distance is consistent with the physical meaning represented by the pseudo range obtained by tracking of the receiver, and the pseudo range difference obtained by subtracting the troposphere error from the ionosphere error is suitable for being used as the observed quantity of Kalman filtering. Because the system does not adjust the clock drift, the clock drift is changed according to the crystal oscillator attribute, the clock drift estimated by the receiver and the clock drift estimated by the tight combination do not conflict, and therefore, the estimated value of the clock drift of the tight combination does not need to be reset.
Specifically, the tightly combined clock difference model includes a dynamic model of a tightly combined clock difference equivalent distance error and a clock drift equivalent speed error, and in the present scheme, the dynamic model of the tightly combined clock difference equivalent distance error and the clock drift equivalent speed error is as follows:
wherein, δ tru＝ctr，trIs the receiver clock drift, c is the speed of light,
is the driving white noise of the receiver clock error equivalent range error,
is the driving white noise of the receiver clock drift equivalent speed error.
The tightly combined clock difference model is substantially identical to the clock difference model of the guard receiver 110. In this embodiment, the clock difference model is also used, which aims to be compatible with receiver positioning and tight combination positioning, not only retains the advantages of tight combination, but also retains the original functions and performance of the receiver, and the clock difference adjusted by each epoch of the receiver is used to reset the clock difference of the tight combination, i.e., to correct the initial state of the tight combination clock difference model.
In one embodiment, before the inertial navigation system 120 is configured to obtain position, velocity, and attitude information of the inertial navigation system through strapdown solution and calculate pseudorange estimates and pseudorange rate estimates in combination with the position and velocity of the received satellite, the method further includes: the inertial navigation system 120 is used for initial setup and initial alignment of the inertial navigation system according to the received position and velocity of the satellite receiver 110.
Specifically, usually, according to different application scenarios, an alignment method and a criterion for determining whether alignment is completed or not are selected, in this embodiment, a dynamic alignment method (in a vehicle-mounted environment) is adopted, that is, when the satellite navigation system continuously positions and the carrier moving speed exceeds a preset threshold, the carrier moving course is calculated by using the speed output by the satellite navigation system, so as to effectively improve the accuracy, and the initial alignment attitude is approximately given by the following formula:
γ＝0
θ＝0
wherein, gamma, theta,
Respectively representing roll angle, pitch angle and course angle, vN、vENorth and east velocities in the northeast coordinate system, respectively.
The satellite inertial tightly combined clock difference correction system is compatible with positioning of the satellite navigation receiver 110 and satellite inertial tightly combined positioning, the output result of the system can be the tightly combined result, and can also be the receiver result, the tightly combined clock difference can be corrected by the receiver clock difference, the receiver adjusts the clock difference in real time and corrects a model in the tight combination realization, the requirement on the crystal oscillator precision of the receiver can be reduced, the tightly combined clock difference is corrected in each epoch, the navigation accuracy is effectively improved, the tightly combined navigation system supports real-time navigation data processing and data post-processing, and the application scene of the navigation system is expanded.
In one embodiment, as shown in fig. 2, a satellite inertial force combined clock difference correction method includes the following steps:
step S110: and the satellite signal is received by the satellite guide receiver and is processed to obtain a pseudo range, a pseudo range rate and a position and a speed of the satellite between the satellite and the satellite guide receiver.
Step S120: and the satellite navigation receiver carries out PVT calculation according to the pseudo range and the pseudo range rate to obtain the position, the speed, the clock difference and the clock drift of the satellite navigation receiver, and adjusts the current clock difference according to the clock difference and the clock drift to obtain the actually adjusted clock difference. In this embodiment, after step S110 and before step S120, the method further includes: acquiring the number of connected satellites, and returning to the step S120 when the number of the connected satellites is greater than or equal to a preset threshold; and returning to the step S110 when the number of satellites is less than the preset threshold.
Step S130: and the satellite navigation receiver sends the pseudo range, the pseudo range rate and the actually adjusted clock difference to the satellite inertial tight combination clock difference correction device, and sends the position and the speed of the satellite and the position and the speed of the satellite navigation receiver to the inertial navigation system.
Step S140: the inertial navigation system obtains position, speed and attitude information of the inertial navigation system through strapdown resolving, and calculates a pseudo-range estimation value and a pseudo-range rate estimation value by combining the received position and speed of the satellite. In this embodiment, step S140 includes: and the inertial navigation system carries out initial setting and initial alignment on the inertial navigation system according to the received position and speed of the satellite navigation receiver.
Step S150: and the inertial navigation system sends the pseudo-range estimated value, the pseudo-range rate estimated value and position, speed and attitude information of the inertial navigation system to the inertial tight combination clock error correction device.
Step S160: and the satellite-inertial tight combination clock difference correction device performs tight combination clock difference correction on the position, speed and attitude information of the inertial navigation system according to the received pseudo range, pseudo range rate, actual adjustment clock difference, pseudo range estimation value and pseudo range rate estimation value, obtains and outputs the corrected position, speed and attitude information of the inertial navigation system. In the present embodiment, step S160 includes step 162 and step 164.
Step 162: and performing close combination filtering according to the pseudo range, the pseudo range rate, the actual adjustment clock error, the pseudo range estimation value and the pseudo range rate estimation value, and calculating to obtain the error of the inertial navigation system. In this embodiment, step 162 includes steps 1622 through 1626.
Step 1622: and acquiring the clock error of the current time predicted by the clock error model.
Step 1624: and calculating the predicted value of the clock difference at the current time according to the clock difference at the current time predicted by the clock difference model and the actually adjusted clock difference.
Step 1626: and performing close combination filtering according to the pseudo range, the pseudo range rate, the pseudo range estimation value and the pseudo range rate estimation value of the clock difference predicted value at the current moment, and calculating to obtain the error of the inertial navigation system.
Step 164: and correcting the position, speed and attitude information of the inertial navigation system according to the error of the inertial navigation system to obtain and output the position, speed and attitude information corrected by the inertial navigation system.
In a more detailed embodiment, as shown in fig. 3, the satellite signal is received by the satellite guidance receiver, and the received satellite signal is processed to calculate a pseudorange and a pseudorange rate of the satellite; the pseudo range and the pseudo range rate of the satellite are used as observed quantities, PVT resolving is carried out according to the relation between the observed quantities and the position, the speed, the clock error and the clock drift of the receiver, namely when the number of the satellites is more than or equal to 4, the position, the speed, the clock error and the clock drift of the satellite receiver are obtained through calculation by adopting a (weighted) least square method or a filtering method; correspondingly advancing or delaying local time according to the clock difference and the clock drift obtained by calculation, namely adjusting the clock difference of the receiver to generate a corresponding local pseudo satellite signal so as to finish continuous tracking of the satellite signal; judging whether the inertial navigation system is aligned, if so, directly carrying out strapdown calculation, if not, carrying out initial setting on the inertial navigation system according to the position and the speed of the satellite navigation receiver, then carrying out initial alignment and strapdown calculation on the inertial navigation system, and calculating to obtain the position, the speed and the attitude information of the carrier; in the integrated navigation part, a Kalman filtering algorithm is adopted, and in the implementation process, a state equation of the system is established according to an error propagation model of inertial navigation and a dynamic model of a clock-error clock drift, wherein the dynamic model of the clock-error clock drift can usually adopt a random walk model or a first-order Markov model; and establishing an observation equation according to the relation between the observed quantity and the state quantity. Generally, the resolving frequency of inertial navigation is higher than that of satellite navigation PVT, a satellite navigation PVT resolving period is used as a system filtering estimation period in the filtering process, and the inertial navigation resolving period is used as a system filtering prediction period. And (3) subtracting a pseudo range and a pseudo range rate between the satellite and the satellite tracked by the satellite guide receiver in each epoch from a pseudo range estimation value and a pseudo range rate estimation value estimated by inertial navigation estimation to serve as an observed quantity of tight combination filtering, predicting the clock difference of the current time according to a clock difference clock drift model (such as a random walk model or a first-order Markov model) by tight combination, and subtracting the clock difference actually adjusted by the last epoch satellite guide receiver to obtain a clock difference predicted value of the tight combination current time. At the moment, the distance between the satellite calculated by inertial navigation and the satellite navigation receiver is added with troposphere error, ionosphere error and clock error predicted value, the distance is consistent with the physical meaning represented by pseudo-range obtained by tracking of the satellite navigation receiver, and pseudo-range difference obtained by subtracting the troposphere error from the ionosphere error is suitable for being used as the observed quantity of Kalman filtering; after the tight combination filtering is finished, estimating to obtain the inertial navigation error including a position error, a speed error and an attitude error, feeding back and correcting the position, the speed and the attitude of the inertial navigation and outputting the inertial navigation error.
The satellite inertial tight combination clock error correction method is compatible with satellite navigation receiver positioning and satellite inertial tight combination positioning, the output result can be used for correcting the satellite navigation tight combination clock error through the receiver clock error, the receiver adjusts the clock error in real time and corrects a model in tight combination realization, the requirement on the crystal oscillator precision of the receiver can be reduced, the satellite navigation error is corrected in each epoch, the navigation accuracy is effectively improved, real-time navigation data processing and data post-processing are supported, and the application scene of a navigation system is expanded.
The technical features of the embodiments described above may be arbitrarily combined, and for the sake of brevity, all possible combinations of the technical features in the embodiments described above are not described, but should be considered as being within the scope of the present specification as long as there is no contradiction between the combinations of the technical features.
The above-mentioned embodiments only express several embodiments of the present invention, and the description thereof is more specific and detailed, but not construed as limiting the scope of the invention. It should be noted that, for a person skilled in the art, several variations and modifications can be made without departing from the inventive concept, which falls within the scope of the present invention. Therefore, the protection scope of the present patent shall be subject to the appended claims.
Claims (10)
1. A satellite inertial compression combined clock difference correction system is characterized by comprising a satellite navigation receiver, an inertial navigation system and a satellite inertial compression combined clock difference correction device, wherein the satellite navigation receiver is connected with the inertial navigation system and the satellite inertial compression combined clock difference correction device, the inertial navigation system is connected with the satellite inertial compression combined clock difference correction device,
the satellite guide receiver is used for receiving satellite signals and processing the satellite signals to obtain pseudo ranges, pseudo range rates and positions and speeds of satellites between the satellites and the satellite guide receiver; performing PVT (physical vapor transport) calculation according to the pseudo range and the pseudo range rate to obtain the position, the speed, the clock error and the clock drift of the guide receiver; adjusting the current clock difference according to the clock difference and the clock drift to obtain an actual adjusted clock difference, and sending the pseudo range, the pseudo range rate and the actual adjusted clock difference to the inertial measurement tight combination clock difference correction device; sending the position and the speed of the satellite and the position and the speed of the satellite navigation receiver to the inertial navigation system;
the inertial navigation system is used for obtaining position, speed and attitude information of the inertial navigation system through strapdown resolving, calculating to obtain a pseudo-range estimation value and a pseudo-range rate estimation value by combining the received position and speed of the satellite, and sending the pseudo-range estimation value, the pseudo-range rate estimation value and the position, speed and attitude information of the inertial navigation system to the satellite inertial navigation tight combination clock error correction device;
and the satellite-inertial tight combination clock difference correction device is used for performing tight combination clock difference correction on the position, speed and attitude information of the inertial navigation system according to the received pseudo range, the pseudo range rate, the actual adjustment clock difference, the pseudo range estimation value and the pseudo range rate estimation value to obtain and output the position, speed and attitude information after the inertial navigation system is corrected.
2. The system of claim 1, wherein the device for correcting the inertial tight combined clock difference is configured to perform the close combined clock difference correction on the position, velocity and attitude information of the inertial navigation system according to the received pseudorange, pseudorange rate, actually adjusted clock difference, pseudorange estimation value and pseudorange rate estimation value, obtain and output the corrected position, velocity and attitude information of the inertial navigation system, and includes:
the satellite-inertial tight combination clock error correction device is used for carrying out tight combination filtering according to the pseudo range, the pseudo range rate, the actual adjustment clock error, the pseudo range estimation value and the pseudo range rate estimation value, and calculating to obtain an error of the inertial navigation system;
and the satellite-inertial tight combination clock error correction device is used for correcting the position, speed and attitude information of the inertial navigation system according to the error of the inertial navigation system to obtain and output the position, speed and attitude information corrected by the inertial navigation system.
3. The system of claim 2, wherein the means for correcting the closed-loop inertial measurement system clock error is configured to perform close-loop filtering based on the pseudorange, the pseudorange rate, the actual adjusted clock error, the pseudorange estimate, and the pseudorange rate estimate to calculate an error of the inertial navigation system, and comprises:
the satellite inertial force combined clock difference correction device is used for acquiring the clock difference of the current moment predicted by the clock difference model;
the satellite inertial force combined clock difference correction device is used for calculating a predicted value of the clock difference at the current time according to the clock difference at the current time predicted by the clock difference model and the actually adjusted clock difference;
and the satellite-inertial tight combination clock difference correction device is used for performing tight combination filtering according to the clock difference predicted value at the current moment, the pseudo range rate, the pseudo range estimated value and the pseudo range rate estimated value, and calculating to obtain the error of the inertial navigation system.
4. The system of claim 1, wherein the satellite navigation receiver is configured to receive satellite signals, process the satellite signals to obtain pseudoranges, pseudorange rates, and positions and velocities of satellites, and perform PVT solution according to the pseudoranges and the pseudorange rates to obtain the positions, velocities, clock biases and clock drifts of the satellite navigation receiver, and the system comprises:
the satellite navigation receiver is used for acquiring the number of connected satellites, and when the number of the satellites is larger than or equal to a preset threshold value, PVT resolving is carried out according to the pseudo range and the pseudo range rate to obtain the position, the speed, the clock error and the clock drift of the satellite navigation receiver;
and returning to receive satellite signals when the number of the satellites is smaller than a preset threshold, and processing the satellite signals to obtain pseudo-range, pseudo-range rate, and position and speed of the satellites.
5. The system of claim 1, wherein the inertial navigation system is configured to obtain position, velocity, and attitude information of the inertial navigation system through strapdown calculation, and before calculating pseudorange estimates and pseudorange rate estimates according to the received positions and velocities of the satellites, further comprising:
and the inertial navigation system is used for carrying out initial setting and initial alignment on the inertial navigation system according to the received position and speed of the satellite navigation receiver.
6. A satellite inertial force combined clock difference correction method is characterized by comprising the following steps:
the satellite signal is received by the satellite guide receiver and processed to obtain a pseudo range, a pseudo range rate and a position and a speed of the satellite between the satellite and the satellite guide receiver;
the satellite navigation receiver carries out PVT resolving according to the pseudo range and the pseudo range rate to obtain the position, the speed, the clock error and the clock drift of the satellite navigation receiver, and the current clock error is adjusted according to the clock error and the clock drift to obtain an actual adjusted clock error;
the satellite navigation receiver sends the pseudo range, the pseudo range rate and the actually adjusted clock difference to a satellite inertial tight combination clock difference correction device, and sends the position and the speed of the satellite and the position and the speed of the satellite navigation receiver to an inertial navigation system;
the inertial navigation system obtains position, speed and attitude information of the inertial navigation system through strapdown resolving, and calculates a pseudo-range estimation value and a pseudo-range rate estimation value by combining the received position and speed of the satellite;
the inertial navigation system sends the pseudo-range estimated value, the pseudo-range rate estimated value and position, speed and attitude information of the inertial navigation system to the satellite-inertial tight combination clock difference correction device;
and the satellite-inertial tight combination clock difference correction device performs tight combination clock difference correction on the position, speed and attitude information of the inertial navigation system according to the received pseudo range, the pseudo range rate, the actual adjustment clock difference, the pseudo range estimation value and the pseudo range rate estimation value, so as to obtain and output the position, speed and attitude information corrected by the inertial navigation system.
7. The method for correcting the inertial tight combination clock difference according to claim 6, wherein the device for correcting the inertial tight combination clock difference performs the tight combination clock difference correction on the position, the velocity and the attitude information of the inertial navigation system according to the received pseudo range, the received pseudo range rate, the received actually adjusted clock difference, the received pseudo range estimation value and the received pseudo range rate estimation value, so as to obtain and output the position, the velocity and the attitude information after the inertial navigation system is corrected, and the method comprises:
performing close combination filtering according to the pseudo range, the pseudo range rate, the actual adjustment clock error, the pseudo range estimation value and the pseudo range rate estimation value, and calculating to obtain an error of the inertial navigation system;
and correcting the position, speed and attitude information of the inertial navigation system according to the error of the inertial navigation system to obtain and output the position, speed and attitude information corrected by the inertial navigation system.
8. The method for correcting inertial tightly-combined clock difference according to claim 7, wherein the step of performing tight-combined filtering by the device for correcting inertial tightly-combined clock difference according to the pseudorange, the pseudorange rate, the actually adjusted clock difference, the pseudorange estimation value, and the pseudorange rate estimation value to calculate an error of the inertial navigation system includes:
acquiring the clock error of the current time predicted by the clock error model;
calculating a predicted value of the clock difference at the current time according to the clock difference at the current time predicted by the clock difference model and the actual adjustment clock difference;
and performing close combination filtering according to the current clock difference predicted value, the pseudo range rate, the pseudo range estimated value and the pseudo range rate estimated value, and calculating to obtain an error of the inertial navigation system.
9. The method for correcting the inertial force combined clock offset according to claim 6, wherein after the satellite signal is received by the satellite receiver and processed to obtain a pseudorange and a pseudorange rate, PVT solution is performed according to the pseudorange and the pseudorange rate, and before the position, the speed, the clock offset and the clock drift of the satellite receiver are obtained, the method comprises:
acquiring the number of connected satellites, and when the number of the satellites is larger than or equal to a preset threshold value, performing PVT (virtual reality) calculation according to the pseudo range and the pseudo range rate to obtain the position, the speed, the clock error and the clock drift of the satellite navigation receiver;
and returning to the step of receiving satellite signals and processing the satellite signals to obtain the pseudo range and the pseudo range rate when the number of the satellites is smaller than a preset threshold value.
10. The method for correcting inertial force combination clock difference according to claim 6, wherein before the inertial navigation system obtains position, velocity and attitude information of the inertial navigation system through strapdown calculation and estimates pseudorange estimated values and pseudorange rate estimated values according to the received position and velocity of the satellite, the method further comprises:
and the inertial navigation system carries out initial setting and initial alignment on the inertial navigation system according to the received position and speed of the satellite navigation receiver.
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| CN112394377A (en) * | 2019-08-14 | 2021-02-23 | Oppo广东移动通信有限公司 | Navigation method, navigation device, electronic equipment and storage medium |
| CN111595331A (en) * | 2019-12-10 | 2020-08-28 | 上海航天控制技术研究所 | Clock model assisted inertial/satellite/relative ranging information combined navigation method |
| CN111256691A (en) * | 2020-02-17 | 2020-06-09 | 苏州芯智谷智能科技有限公司 | Networking hardware time reference establishing method based on GNSS/MEMS inertia combined chip |
| CN112083465A (en) * | 2020-09-18 | 2020-12-15 | 德明通讯（上海）有限责任公司 | Position information acquisition system and method |
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